Film cooling hole for a turbine airfoil
A film cooling hole for a turbine airfoil. the first embodiment is a film cooling hole aligned with the stream-wise direction of the hot gas flow over the hole and includes a metering section followed by a diffuser section. The diffuser section includes a left side wall and a right side wall both having a curvature facing outward and at about the same radius of curvature. The diffuser section also includes a top side wall and a bottom side wall which both have a curvature in the stream-wise direction, and in which the radius of curvature of the top side wall is greater than the radius of curvature of the bottom side wall. A second embodiment is a film cooling hole offset from the stream-wise direction of flow and includes the four walls with a curvature, but in which the left side wall has a greater radius of curvature than the right side wall due to the stream-wise offset of the hole.
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1. Field of the Invention
The present invention relates generally to air cooled turbine airfoils, and more specifically to a film cooling hole for the airfoils.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
In a gas turbine engine, a turbine comprises a number of stages of stator vanes and rotor blades used to convert the energy from a hot gas flow into mechanical energy used to drive the rotor shaft. The efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine. However, the highest temperature allowable is dependent on the material properties of the first stage airfoils (vanes and blades) and the amount of cooling provided. Once the material properties have been established, higher temperatures can be used if adequate cooling of the airfoils is provided.
Current airfoil cooling designs make use of internal convection and impingement cooling, and film cooling of the external airfoil surfaces that are exposed to the high temperature gas flow. Film cooling provides a blanket of cooling air over the airfoil surface that—in theory—prevents the hot gas flow from making contact with the airfoil surface. One major objective of a turbine airfoil designer is to maximize the effect of the cooling air while minimizing the usage of the cooling air in order to increase the efficiency of the engine, since the pressurized cooling air used for cooling the airfoils is bled of from the compressor of the engine. The bled off cooling air becomes wasted work.
Prior art film holes pass straight through the airfoil wall at a constant diameter and exit at an angle to the airfoil surface. Some of the cooling air is consequently ejected directly into the mainstream hot gas flow and causing turbulence, coolant dilution and a loss of downstream film effectiveness. Also, the hole breakout in the stream-wise elliptical shape will induce stress problems in a blade application.
The above described problems associated with turbine airfoil film cooling holes can be reduced by incorporating the film cooling hole geometry of the present invention into the prior art airfoil cooling design. The film hole of the present invention includes a curved diffusion hole in which each individual inner wall of the film hole is constructed with a various radius of curvature independent to each other. The unique film cooling hole design will allow for radial diffusion of the stream-wise oriented flow which combines the best aspects of both radial and stream-wise straight holes.
In one embodiment, the film hole is aligned with the stream-wise direction of the hot gas flow and the sides walls of the film hole have about the same amount of curvature. In a second embodiment, the film hole has side walls at different amounts of curvature to form a compound angle in which the stream-wise direction is not parallel to the film hole axis.
The film cooling hole of the present invention is for use in an air cooled turbine airfoil such as a rotor blade or a stator vane of a gas turbine engine such as an industrial gas turbine (IGT) engine. However, the film cooling hole can be used in other devices in which film cooling of a surface is required in order to protect the surface from the effects of a high temperature gas flow passing over the surface. A combustor in a power plant or in a gas turbine engine requires film cooling and can make use of the film cooling hole of the present invention.
The first embodiment of the film cooling hole of the present invention is shown in
In the first embodiment film hole 30, the side walls 33 and 34 have about the same radius of curvature (R3=R4) while the top side wall 35 has a radius of curvature R1 greater than the bottom side wall 36 radius of curvature R2.
A second embodiment of the film cooling hole 40 is shown in
The film hole 40 includes a top side wall 45 and a bottom side wall 46 as shown in
In the stream-wise direction, the curved wall at the upstream (35 in
In the spanwise direction, the radial outward and radial inward film cooling hole walls (33 and 34 in
In summary, the various radius of curvature diffusion film hole has the expansion radial and rearward hole surfaces curved toward both the airfoil trailing edge and spanwise directions. Coolant penetration into the gas path is thus minimized, yielding good buildup of the coolant sub-boundary layer next to the airfoil surface, lower aerodynamic mixing losses due to low angle of cooling air injection, better film coverage in the spanwise direction and high film effectiveness for a longer distance downstream of the film hole. The end result of both benefits produces a better film cooling effectiveness level for the turbine airfoil.
Claims
1. A film cooling hole for use in an air cooled turbine airfoil, the film cooling hole comprising:
- a metering section forming an inlet to the film cooling hole;
- a diffuser section downstream from the metering section;
- the diffuser section being formed with a left side wall and a right side wall, and a top side wall and a bottom side wall;
- the left side wall and the right side wall both having a radius of curvature in an outward direction; and,
- the top side wall and the bottom side wall both having a radius of curvature toward a stream-wise direction of the hot gas flow.
2. The film cooling hole of claim 1, and further comprising:
- the metering hole axis is aligned with the stream-wise direction of the hot gas flow; and,
- the radius of curvature of the left side wall and the right side wall is substantially equal.
3. The film cooling hole of claim 2, and further comprising:
- the radius of curvature of the top side wall is greater than the radius of curvature of the bottom side wall.
4. The film cooling hole of claim 3, and further comprising:
- the left side wall and the right side wall are offset from the metering hole axis in the range of from 7 degrees to 15 degrees.
5. The film cooling hole of claim 4, and further comprising:
- the top side wall is offset from the metering hole axis from zero to 5 degrees; and,
- the bottom side wall is offset from the metering hole axis from 15 to 25 degrees.
6. The film cooling hole of claim 1, and further comprising:
- the metering hole axis is significantly offset from the stream-wise direction of the hot gas flow; and,
- the radius of curvature of the left side wall is greater than the radius of curvature of the right side wall.
7. The film cooling hole of claim 6, and further comprising:
- the radius of curvature of the top side wall is greater than the radius of curvature of the bottom side wall.
8. The film cooling hole of claim 7, and further comprising:
- the left side wall is offset from the metering hole axis from zero to 5 degrees; and,
- the right side wall is offset from the metering hole axis from 15 degrees to 25 degrees.
9. The film cooling hole of claim 8, and further comprising:
- the top side wall is offset from the metering hole axis from zero to 5 degrees; and,
- the bottom side wall is offset from the metering hole axis from 15 to 25 degrees.
10. A turbine airfoil for use in a gas turbine engine, the airfoil comprising:
- a pressure side wall and a suction side wall defining the airfoil surface;
- an internal cooling circuit to provide cooling for the airfoil; and,
- a plurality of film cooling holes connected to the internal cooling circuit, the film cooling holes further comprising a metering section and a diffuser section, the diffuser section including a left side wall and a right side wall both with a curvature facing outward, the diffuser section including a top side wall and a bottom side wall both with a curvature facing toward the bottom of the diffuser.
11. The turbine airfoil of claim 10, and further comprising:
- the metering hole axis is aligned with the stream-wise direction of the hot gas flow; and,
- the radius of curvature of the left side wall and the right side wall is substantially equal.
12. The film cooling hole of claim 11, and further comprising:
- the radius of curvature of the top side wall is greater than the radius of curvature of the bottom side wall.
13. The film cooling hole of claim 12, and further comprising:
- the left side wall and the right side wall are offset from the metering hole axis in the range of from 7 degrees to 15 degrees.
14. The film cooling hole of claim 13, and further comprising:
- the top side wall is offset from the metering hole axis from zero to 5 degrees; and,
- the bottom side wall is offset from the metering hole axis from 15 to 25 degrees.
15. The film cooling hole of claim 10, and further comprising:
- the metering hole axis is significantly offset from the stream-wise direction of the hot gas flow; and,
- the radius of curvature of the left side wall is greater than the radius of curvature of the right side wall.
16. The film cooling hole of claim 15, and further comprising:
- the radius of curvature of the top side wall is greater than the radius of curvature of the bottom side wall.
17. The film cooling hole of claim 16, and further comprising:
- the left side wall is offset from the metering hole axis from zero to 5 degrees; and,
- the right side wall is offset from the metering hole axis from 15 degrees to 25 degrees.
18. The film cooling hole of claim 17, and further comprising:
- the top side wall is offset from the metering hole axis from zero to 5 degrees; and,
- the bottom side wall is offset from the metering hole axis from 15 to 25 degrees.
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Type: Grant
Filed: Nov 19, 2007
Date of Patent: Nov 29, 2011
Assignee: Florida Turbine Technologies, Inc. (Jupiter, FL)
Inventor: George Liang (Palm City, FL)
Primary Examiner: Igor Kershteyn
Attorney: John Ryznic
Application Number: 11/986,033
International Classification: F01D 5/08 (20060101);